US20140023830A1 - Method of manufacturing a structure - Google Patents
Method of manufacturing a structure Download PDFInfo
- Publication number
- US20140023830A1 US20140023830A1 US13/944,489 US201313944489A US2014023830A1 US 20140023830 A1 US20140023830 A1 US 20140023830A1 US 201313944489 A US201313944489 A US 201313944489A US 2014023830 A1 US2014023830 A1 US 2014023830A1
- Authority
- US
- United States
- Prior art keywords
- stack
- thickness
- thickness control
- ramp
- plies
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
- 238000004519 manufacturing process Methods 0.000 title claims abstract description 12
- 239000002131 composite material Substances 0.000 claims abstract description 51
- 238000000034 method Methods 0.000 claims abstract description 19
- 230000007423 decrease Effects 0.000 claims abstract description 13
- 238000005452 bending Methods 0.000 claims description 4
- 238000003754 machining Methods 0.000 description 6
- 229920002430 Fibre-reinforced plastic Polymers 0.000 description 3
- OKTJSMMVPCPJKN-UHFFFAOYSA-N Carbon Chemical compound [C] OKTJSMMVPCPJKN-UHFFFAOYSA-N 0.000 description 2
- 229910052799 carbon Inorganic materials 0.000 description 2
- 230000000694 effects Effects 0.000 description 2
- 239000011151 fibre-reinforced plastic Substances 0.000 description 2
- 210000001503 joint Anatomy 0.000 description 2
- UOORRWUZONOOLO-OWOJBTEDSA-N (E)-1,3-dichloropropene Chemical compound ClC\C=C\Cl UOORRWUZONOOLO-OWOJBTEDSA-N 0.000 description 1
- 230000000712 assembly Effects 0.000 description 1
- 238000000429 assembly Methods 0.000 description 1
- 230000001419 dependent effect Effects 0.000 description 1
- 239000003365 glass fiber Substances 0.000 description 1
- 238000010438 heat treatment Methods 0.000 description 1
- 238000009434 installation Methods 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
Images
Classifications
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B32—LAYERED PRODUCTS
- B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
- B32B3/00—Layered products comprising a layer with external or internal discontinuities or unevennesses, or a layer of non-planar shape; Layered products comprising a layer having particular features of form
- B32B3/26—Layered products comprising a layer with external or internal discontinuities or unevennesses, or a layer of non-planar shape; Layered products comprising a layer having particular features of form characterised by a particular shape of the outline of the cross-section of a continuous layer; characterised by a layer with cavities or internal voids ; characterised by an apertured layer
- B32B3/263—Layered products comprising a layer with external or internal discontinuities or unevennesses, or a layer of non-planar shape; Layered products comprising a layer having particular features of form characterised by a particular shape of the outline of the cross-section of a continuous layer; characterised by a layer with cavities or internal voids ; characterised by an apertured layer characterised by a layer having non-uniform thickness
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C3/00—Wings
- B64C3/18—Spars; Ribs; Stringers
- B64C3/185—Spars
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C3/00—Wings
- B64C3/26—Construction, shape, or attachment of separate skins, e.g. panels
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29C—SHAPING OR JOINING OF PLASTICS; SHAPING OF MATERIAL IN A PLASTIC STATE, NOT OTHERWISE PROVIDED FOR; AFTER-TREATMENT OF THE SHAPED PRODUCTS, e.g. REPAIRING
- B29C70/00—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts
- B29C70/04—Shaping composites, i.e. plastics material comprising reinforcements, fillers or preformed parts, e.g. inserts comprising reinforcements only, e.g. self-reinforcing plastics
- B29C70/28—Shaping operations therefor
- B29C70/30—Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core
- B29C70/34—Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core and shaping or impregnating by compression, i.e. combined with compressing after the lay-up operation
- B29C70/342—Shaping by lay-up, i.e. applying fibres, tape or broadsheet on a mould, former or core; Shaping by spray-up, i.e. spraying of fibres on a mould, former or core and shaping or impregnating by compression, i.e. combined with compressing after the lay-up operation using isostatic pressure
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29K—INDEXING SCHEME ASSOCIATED WITH SUBCLASSES B29B, B29C OR B29D, RELATING TO MOULDING MATERIALS OR TO MATERIALS FOR MOULDS, REINFORCEMENTS, FILLERS OR PREFORMED PARTS, e.g. INSERTS
- B29K2707/00—Use of elements other than metals for preformed parts, e.g. for inserts
- B29K2707/04—Carbon
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29K—INDEXING SCHEME ASSOCIATED WITH SUBCLASSES B29B, B29C OR B29D, RELATING TO MOULDING MATERIALS OR TO MATERIALS FOR MOULDS, REINFORCEMENTS, FILLERS OR PREFORMED PARTS, e.g. INSERTS
- B29K2709/00—Use of inorganic materials not provided for in groups B29K2703/00 - B29K2707/00, for preformed parts, e.g. for inserts
- B29K2709/08—Glass
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B29—WORKING OF PLASTICS; WORKING OF SUBSTANCES IN A PLASTIC STATE IN GENERAL
- B29L—INDEXING SCHEME ASSOCIATED WITH SUBCLASS B29C, RELATING TO PARTICULAR ARTICLES
- B29L2031/00—Other particular articles
- B29L2031/30—Vehicles, e.g. ships or aircraft, or body parts thereof
- B29L2031/3076—Aircrafts
- B29L2031/3085—Wings
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
- B64C1/06—Frames; Stringers; Longerons ; Fuselage sections
- B64C1/12—Construction or attachment of skin panels
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
- B64C1/26—Attaching the wing or tail units or stabilising surfaces
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/40—Weight reduction
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
- Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
- Y10T428/00—Stock material or miscellaneous articles
- Y10T428/24—Structurally defined web or sheet [e.g., overall dimension, etc.]
- Y10T428/24479—Structurally defined web or sheet [e.g., overall dimension, etc.] including variation in thickness
- Y10T428/24612—Composite web or sheet
Definitions
- the present invention relates to a method of manufacturing a structure in which a laminar composite panel is attached to a first component, and the laminar composite panel is attached to a further component after it has been attached to the first component.
- the panel may be a cover of an aircraft wing
- the first component may be a spar of the wing
- the further component may be a centre wing box.
- FIG. 1 is a schematic cross sectional view of an aircraft wing root joint comprising an aircraft wing box 1 attached to a centre wing box 10 .
- the centre wing box 10 comprises a rib 11 connecting upper and lower root joint fittings 2 , 3 .
- the aircraft wing box 1 comprises upper and lower laminar composite covers 4 , 5 connected by a spar 6 which extends between them.
- the covers 4 , 5 have inner and outer mould lines 4 a, 5 a, 4 b, 5 b.
- the outer mould lines 4 b, 5 b form external aerodynamic surfaces of the wing box 1
- the inner mould lines 4 a, 5 a form internal surfaces of the wing box 1 opposite the outer mould lines 4 b, 5 b.
- the spar 6 is shown in longitudinal section in FIG. 1 .
- the spar 6 is C-shaped in transverse section with upper and lower spar flanges 7 , 8 which extend from a web 9 and engage the inner mould lines 4 a, 5 a of the covers 4 , 5 .
- the spar flanges 7 , 8 terminate outboard of the inboard edge of the spar web 9 to leave spar flange cutback areas where the covers 4 , 5 overhang the flanges 7 , 8 in an inboard span-wise direction towards the root joint fittings 2 , 3 .
- a problem with manufacturing such a joint is that the thickness of the laminar composite covers 4 , 5 cannot be controlled accurately due to the method of manufacturing the laminate and the properties of the composite material. As a result the covers 4 , 5 may not precisely meet the root joint fittings 2 , 3 resulting in either a gap or a clash.
- This problem can be solved by bending the covers 4 , 5 (where they overhang the spar flanges 7 , 8 ) either towards the fittings (to close a gap) or away from the fittings (to avoid a clash). However this will induce stress in the covers 4 , 5 and as a result the covers must be made thicker and heavier to cope with this stress. Also it may be difficult to bend the covers sufficiently, particularly near the cover/spar interface shown in FIG. 1 .
- the thickness of the upper cover 4 (i.e. the distance between the inner mould line 4 a and the outer mould line 4 b ) is particularly critical as it is its outer mould line 4 b (rather than its inner mould line 4 a ) which engages the root joint fitting 2 .
- any increase or decrease in the thickness of the upper cover 4 will affect the alignment between the upper cover 4 and the upper root joint fitting 2 and/or the alignment between the lower cover 5 and the lower root joint fitting 3 .
- the thickness of the lower cover does not affect the alignment between the lower root joint fitting 3 and the lower cover 5 , nor does the thickness of the lower cover affect the alignment between the upper cover 4 and the upper root joint fitting 2 .
- the thickness of the lower cover is critical for the attachment of a buttstrap 12 which connects the lower cover OML 5 b to a lower cover 13 of the centre wing box 10 . Therefore, any thickness variation in the manufacturing process of both covers must be carefully managed.
- a first aspect of the invention provides a method of manufacturing a structure, the method comprising
- steps a), d) and e) may be performed at the same time, but preferably step e) is performed after step d).
- a second aspect of the invention provides a structure comprising:
- the laminar composite panel may be attached at or near its edge to the further component by an overlapping joint, a butt joint, or any other suitable connection.
- a longest one of the thickness control plies extends along only part of a length of the laminar composite panel, preferably less than 50% of its length.
- the thickness of the stack of thickness control plies decreases towards the second edge.
- the laminate composite panel In the event that the laminate composite panel is too thick or thin, then it may be bent to form a ramp after step c).
- the ramp may be inclined in the opposite sense to the ramp in the stack of thickness control plies, or in the same sense.
- CFRP Carbon Fibre Reinforced Plastic
- the further component is a centre wing box of an aircraft wing
- the first component is a spar of an aircraft wing
- the panel assembly is an aircraft wing cover assembly.
- the further component may be an aircraft fuselage skin panel
- the first component may be a frame of an aircraft fuselage
- the panel assembly may be an aircraft fuselage skin panel assembly which is attached to the first component by a butt-strap.
- FIG. 1 is a cross-sectional schematic view of an aircraft root wing joint
- FIGS. 2 and 3 illustrate a method of managing the thickness of a composite panel
- FIG. 4 is a cross-sectional schematic view of an aircraft root wing joint where the thicknesses of the wing box covers have been managed using the method illustrated by FIGS. 2 , 3 ;
- FIG. 5 is a perspective view of the wing box shown in FIG. 4 ;
- FIG. 6 is a cross-sectional schematic view of a cover to spar interface where the cover has a thickness at the lower end of a manufacturing tolerance
- FIG. 7 is a similar view to FIG. 6 but where the cover has a thickness at the upper end of a manufacturing tolerance.
- FIG. 2 show a pair of partially cured Carbon Fibre Reinforced Plastic (CFRP) aircraft wing covers 20 , 22 on a mould tool 24 .
- the covers 20 , 22 have outer mould lines (OMLs) 21 , 23 , which engage the mould tool 24 , and non-moulded inner mould lines (IMLs) 25 , 27 opposite the outer mould lines 21 , 23 .
- the covers are moulded by laying a vacuum bag (not shown) over the IMLs, applying a vacuum between the vacuum bag and the covers to press the covers against the mould tool 24 , and heating the covers to cure them.
- the thicknesses of the covers 20 , 22 are not accurately controllable and can range from a lower tolerance value X (see FIG. 2 ) to an upper tolerance value Y (see FIG. 2 ). As explained above with respect to FIG. 1 , these relatively large uncertainties in cover thickness are problematic at the aircraft wing root joint.
- a stack 26 of uncured thickness control plies is laid up on the IMLs of the covers 20 , 22 and the cover/thickness control ply assemblies are either co-cured, or the thickness control plies are co-bonded onto the cover during a further curing cycle.
- the stack 26 has a ramped section 32 with a ramp 33 (that is, an inclined surface) where the thickness of the stack 26 decreases towards its outboard edge, and a flat section 34 towards its inboard edge.
- the ramped section 32 may be formed, for example, by laying up progressively shorter thickness control plies on top of one another or by selectively discontinuing internal plies. Referring to FIGS.
- the thickness of the cured panel assembly is first measured, and then if necessary some of the thickness control plies are removed from the stack 26 by a machining process. The number of plies which are removed will thus depend on the measured thickness.
- the distance D is subject only to the tolerances of the machining and measuring processes, regardless of the thickness of the covers 20 , 22 .
- the tolerances involved in machining and measuring are typically far less than those involved in manufacturing CFRP covers 20 , 22 . Therefore, the distance D can be controlled with far greater accuracy than the thickness of the laminar composite covers 20 , 22 .
- the stack 26 of thickness control plies extends along only a relatively small portion of the span of the wing close to its inboard edge, the span-wise length of the thickness control plies being determined by the stress allowable of the covers. This minimises both the size of the required machining tool and the weight of the added thickness control plies.
- FIG. 4 shows a cross sectional schematic view of the root of an aircraft wing box 39 .
- the wing box comprises upper and lower laminar composite covers 40 , 41 which are bolted to a spar 42 which extends between them. Note that only the inboard root end of the spar and wing box is shown in FIG. 4 .
- the spar 42 comprises upper and lower spar flanges 44 , 45 extending from a spar web 46 , and the flanges 44 , 45 engage the upper and lower covers respectively.
- the spar flanges terminate outboard of the inboard edge of the spar web 46 to leave spar flange cutback areas where the covers overhang the flanges in an inboard span-wise direction.
- the covers 40 , 41 have been manufactured by the process described above with reference to FIGS. 2-3 . As the upper and lower covers (and the upper and lower spar flanges) are similar, only the upper cover (and the upper spar flange) will be described in detail.
- the upper cover 40 has a moulded surface (OML) 48 , a non-moulded surface (IML) 50 opposite the OML 48 and a stack 52 of thickness control plies co-bonded or co-cured to the IML 50 of the cover.
- the thickness of the stack 52 control plies has been controlled as described above to form a controlled IML surface 53 which is a precise distance D from the OML 48 .
- the stack 52 consists of a ramped outboard section 54 with a ramp 55 where the thickness of the stack 52 decreases towards its relatively thin outboard edge, and a flat inboard section 56 which terminates at the same span- wise position as the spar flange 44 .
- the spar web 46 converges at its inboard end, and the flange 44 is joggled to follow the line of the spar web and form a ramped surface 58 and a flat un-ramped surface 59 .
- the ramp 58 in the spar flange engages the ramp 55 in the stack, and the un-ramped surface 59 of the spar flange engages the machined surface 53 of the stack 52 .
- the covers and spar flanges are bolted together.
- the upper and lower covers 40 , 41 are also bolted to respective cruciform and tri-form root joint fittings 65 , 67 of a centre wing box.
- the spar web 46 is attached to a spar web (not shown) of the centre wing box via a vertical buttstrap (also not shown).
- the centre wing box also has a rib 68 (conventionally referred to as “rib one”) connecting the upper and lower root joint fittings 65 , 67 .
- rib one commonly referred to as “rib one”
- the covers 40 , 41 can be made thinner and lighter.
- the buttstrap 69 can be attached with less measuring and machining, thus improving the quality and reducing assembly time.
- the thickness of the cover is equal to its ‘design value’—that is, the cover has a thickness which lies in the centre of a manufacturing tolerance range. Consequently, the outer mould line 48 forms a continuous flat aerodynamic surface, even in the ramped region.
- FIG. 5 is a perspective view of the wing box 39 taken from above with the upper cover removed.
- FIG. 5 also shows ribs 61 , 62 (conventionally referred to as “rib two” and “rib three” respectively) which connect the fore and aft spars of the wing box 39 as well as the upper and lower covers.
- the ribs 61 , 62 are bolted to the upper and lower covers and to the webs 46 of the fore and aft spars.
- FIG. 6 is similar to FIG. 4 and identical features are given the same reference numerals. Note that the lower cover has been omitted for clarity. In this case, the thickness of the cover 40 a is equal to the lower tolerance value X (see FIG. 2 ). In order to achieve the required value of D, the thickness of the stack 52 a of thickness control plies is greater than the thickness of the (machined) stack 52 shown in Figure 4 . When the cover 40 a and the stack 52 a of thickness control plies are bolted to the spar, the portion of the cover outboard of the stack 52 a is bent down towards the upper spar flange 44 . This creates a joggled section 70 a in the cover 40 a which is inclined in the opposite sense to the ramp 55 in the stack of thickness control plies.
- the angle of inclination of the ramp in the joggled section 70 a is typically very shallow (i.e. within the bending allowables) and has little effect on the aerodynamic performance of the wing. Note that the angle of the joggled section 70 a shown in FIG. 6 is exaggerated for clarity.
- FIG. 7 is similar to FIGS. 4 , 6 and identical features will be given the same reference numerals. Note that the lower cover has again been omitted for clarity. In this case the thickness of the cover 40 b is equal to the upper tolerance value Y (illustrated in FIG. 2 ).
- Y illustrated in FIG. 2
- the portion of the cover outboard of the stack 52 b is bent up to create a joggled section 70 b with a ramp which is inclined in the same sense as the ramp in the stack of thickness control plies.
- the ramp in the joggled cover engages the top of the ramp in the spar flange.
- the ramp in the joggled section 70 b is typically very shallow (i.e. within the bending allowables) and has little effect on the aerodynamic performance of the wing. Again, no shimming is required, which saves time and complexity. Note that, again, the angle of the joggled section 70 b shown in FIG. 7 is exaggerated for clarity. Note also that the joggle has a downward inboard incline as opposed to the upward inboard incline of the joggled section 70 a shown in FIG. 6 .
- the stacks 52 , 52 a, 52 b decrease in thickness towards their outboard edges.
- the stacks 52 , 52 a, 52 b may decrease in thickness towards their inboard edges (i.e. the edges proximate the inboard edges of the covers 40 , 40 a, 40 b ).
- the height of the spar web would be matched accordingly with the stacks.
- the thickness of the cured panel assembly is first measured, and then if necessary some of the thickness control plies are removed from the stack 26 by a machining process.
- the thickness of the cured cover may be measured before the thickness control plies are bonded in place, and the number of thickness control plies selected in accordance with the measured thickness (either by peeling off some of the uncured plies in a pre-prepared stack, or by laying up the stack with exactly the right number of plies).
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- Aviation & Aerospace Engineering (AREA)
- Laminated Bodies (AREA)
- Moulding By Coating Moulds (AREA)
- Lining Or Joining Of Plastics Or The Like (AREA)
Abstract
Description
- The present invention relates to a method of manufacturing a structure in which a laminar composite panel is attached to a first component, and the laminar composite panel is attached to a further component after it has been attached to the first component. By way of non-limiting example the panel may be a cover of an aircraft wing, the first component may be a spar of the wing, and the further component may be a centre wing box.
-
FIG. 1 is a schematic cross sectional view of an aircraft wing root joint comprising anaircraft wing box 1 attached to acentre wing box 10. Thecentre wing box 10 comprises arib 11 connecting upper and lowerroot joint fittings aircraft wing box 1 comprises upper and lower laminar composite covers 4, 5 connected by aspar 6 which extends between them. - The
covers outer mould lines outer mould lines wing box 1, while theinner mould lines wing box 1 opposite theouter mould lines - The
spar 6 is shown in longitudinal section inFIG. 1 . Thespar 6 is C-shaped in transverse section with upper andlower spar flanges web 9 and engage theinner mould lines covers spar flanges spar web 9 to leave spar flange cutback areas where the covers 4, 5 overhang theflanges root joint fittings - A problem with manufacturing such a joint is that the thickness of the laminar composite covers 4, 5 cannot be controlled accurately due to the method of manufacturing the laminate and the properties of the composite material. As a result the
covers root joint fittings covers 4, 5 (where they overhang thespar flanges 7, 8) either towards the fittings (to close a gap) or away from the fittings (to avoid a clash). However this will induce stress in thecovers FIG. 1 . - The thickness of the upper cover 4 (i.e. the distance between the
inner mould line 4 a and theouter mould line 4 b) is particularly critical as it is itsouter mould line 4 b (rather than itsinner mould line 4 a) which engages theroot joint fitting 2. Clearly, any increase or decrease in the thickness of theupper cover 4 will affect the alignment between theupper cover 4 and the upper root joint fitting 2 and/or the alignment between thelower cover 5 and the lowerroot joint fitting 3. Conversely, as it is theinner mould line 5 a of thelower cover 5 which engages the lower root joint fitting 3 (rather than itsouter mould line 5 b), the thickness of the lower cover does not affect the alignment between the lowerroot joint fitting 3 and thelower cover 5, nor does the thickness of the lower cover affect the alignment between theupper cover 4 and the upperroot joint fitting 2. However, the thickness of the lower cover is critical for the attachment of abuttstrap 12 which connects thelower cover OML 5 b to alower cover 13 of thecentre wing box 10. Therefore, any thickness variation in the manufacturing process of both covers must be carefully managed. - A first aspect of the invention provides a method of manufacturing a structure, the method comprising
-
- a) forming a panel assembly by bonding a stack of thickness control plies of composite material to a laminar composite panel, the stack of thickness control plies having a first edge proximate an edge of the laminar composite panel, a second edge opposite the first edge, and a ramp where the thickness of the stack of thickness control plies decreases towards the first or second edge;
- b) measuring the thickness of the laminate composite panel or the panel assembly;
- c) controlling the number of thickness control plies in accordance with the measured thickness;
- d) attaching a first component to the panel assembly, the first component having a surface which engages the laminar composite panel and a ramp which engages the ramp in the stack of thickness control plies; and
- e) attaching the laminar composite panel at or near its edge to a further component.
- The various steps of the method may be performed in any order, and indeed some of the steps (for instance steps a), d) and e)) may be performed at the same time, but preferably step e) is performed after step d).
- A second aspect of the invention provides a structure comprising:
-
- a panel assembly comprising a stack of thickness control plies of composite material bonded to a laminar composite panel, the stack of thickness control plies having a first edge proximate an edge of the laminar composite panel, a second edge opposite the first edge, and a ramp where the thickness of the stack of thickness control plies decreases towards the first or second edge;
- a first component attached to the panel assembly, the first component having a surface which engages the laminar composite panel and a ramp which engages the ramp in the stack of thickness control plies; and
- a further component attached to the laminar composite panel at or near its edge.
- The laminar composite panel may be attached at or near its edge to the further component by an overlapping joint, a butt joint, or any other suitable connection.
- Typically a longest one of the thickness control plies extends along only part of a length of the laminar composite panel, preferably less than 50% of its length.
- Preferably the thickness of the stack of thickness control plies decreases towards the second edge.
- In the event that the laminate composite panel is too thick or thin, then it may be bent to form a ramp after step c). The ramp may be inclined in the opposite sense to the ramp in the stack of thickness control plies, or in the same sense.
- Typically at least some of the plies of the laminar composite panel are made from Carbon Fibre Reinforced Plastic (CFRP), although other composite materials may be envisaged such as Glass-Fibre Reinforced Plastic.
- Further preferred features of the present invention are set out in the dependent claims.
- In the embodiment of the invention described the further component is a centre wing box of an aircraft wing, the first component is a spar of an aircraft wing and the panel assembly is an aircraft wing cover assembly. However the invention is applicable to a variety of other joints. For instance the further component may be an aircraft fuselage skin panel, the first component may be a frame of an aircraft fuselage, and the panel assembly may be an aircraft fuselage skin panel assembly which is attached to the first component by a butt-strap.
- Embodiments of the invention will now be described with reference to the accompanying drawings, in which:
-
FIG. 1 is a cross-sectional schematic view of an aircraft root wing joint; -
FIGS. 2 and 3 illustrate a method of managing the thickness of a composite panel; -
FIG. 4 is a cross-sectional schematic view of an aircraft root wing joint where the thicknesses of the wing box covers have been managed using the method illustrated byFIGS. 2 , 3; -
FIG. 5 is a perspective view of the wing box shown inFIG. 4 ; -
FIG. 6 is a cross-sectional schematic view of a cover to spar interface where the cover has a thickness at the lower end of a manufacturing tolerance; and -
FIG. 7 is a similar view toFIG. 6 but where the cover has a thickness at the upper end of a manufacturing tolerance. -
FIG. 2 show a pair of partially cured Carbon Fibre Reinforced Plastic (CFRP) aircraft wing covers 20, 22 on amould tool 24. The covers 20, 22 have outer mould lines (OMLs) 21, 23, which engage themould tool 24, and non-moulded inner mould lines (IMLs) 25, 27 opposite theouter mould lines mould tool 24, and heating the covers to cure them. - The thicknesses of the
covers FIG. 2 ) to an upper tolerance value Y (seeFIG. 2 ). As explained above with respect toFIG. 1 , these relatively large uncertainties in cover thickness are problematic at the aircraft wing root joint. - In order to manage the cover thickness tolerances in the region of the root joint, a
stack 26 of uncured thickness control plies is laid up on the IMLs of thecovers stack 26 has a rampedsection 32 with a ramp 33 (that is, an inclined surface) where the thickness of thestack 26 decreases towards its outboard edge, and aflat section 34 towards its inboard edge. The rampedsection 32 may be formed, for example, by laying up progressively shorter thickness control plies on top of one another or by selectively discontinuing internal plies. Referring toFIGS. 3 a and 3 b, the thickness of the cured panel assembly is first measured, and then if necessary some of the thickness control plies are removed from thestack 26 by a machining process. The number of plies which are removed will thus depend on the measured thickness. This results in machined and curedIML surfaces stack 26 of thickness control plies extends along only a relatively small portion of the span of the wing close to its inboard edge, the span-wise length of the thickness control plies being determined by the stress allowable of the covers. This minimises both the size of the required machining tool and the weight of the added thickness control plies. -
FIG. 4 shows a cross sectional schematic view of the root of anaircraft wing box 39. The wing box comprises upper and lower laminar composite covers 40, 41 which are bolted to aspar 42 which extends between them. Note that only the inboard root end of the spar and wing box is shown inFIG. 4 . Thespar 42 comprises upper andlower spar flanges spar web 46, and theflanges spar web 46 to leave spar flange cutback areas where the covers overhang the flanges in an inboard span-wise direction. Thecovers FIGS. 2-3 . As the upper and lower covers (and the upper and lower spar flanges) are similar, only the upper cover (and the upper spar flange) will be described in detail. - The
upper cover 40 has a moulded surface (OML) 48, a non-moulded surface (IML) 50 opposite theOML 48 and astack 52 of thickness control plies co-bonded or co-cured to theIML 50 of the cover. The thickness of thestack 52 control plies has been controlled as described above to form a controlledIML surface 53 which is a precise distance D from theOML 48. Thestack 52 consists of a rampedoutboard section 54 with aramp 55 where the thickness of thestack 52 decreases towards its relatively thin outboard edge, and a flatinboard section 56 which terminates at the same span- wise position as thespar flange 44. Thespar web 46 converges at its inboard end, and theflange 44 is joggled to follow the line of the spar web and form a rampedsurface 58 and a flatun-ramped surface 59. Theramp 58 in the spar flange engages theramp 55 in the stack, and theun-ramped surface 59 of the spar flange engages the machinedsurface 53 of thestack 52. The covers and spar flanges are bolted together. - Still referring to
FIG. 4 , the upper andlower covers joint fittings spar web 46 is attached to a spar web (not shown) of the centre wing box via a vertical buttstrap (also not shown). - The centre wing box also has a rib 68 (conventionally referred to as “rib one”) connecting the upper and lower root
joint fittings upper cover 40 can be controlled with close accuracy, the problem of gaps/clashes arising between the wing covers 40, 41 and the rootjoint fittings covers buttstrap 69 can be attached with less measuring and machining, thus improving the quality and reducing assembly time. - In the case illustrated in
FIG. 4 (referring to theupper cover 40 but also applicable to the lower cover 41) the thickness of the cover is equal to its ‘design value’—that is, the cover has a thickness which lies in the centre of a manufacturing tolerance range. Consequently, theouter mould line 48 forms a continuous flat aerodynamic surface, even in the ramped region. -
FIG. 5 is a perspective view of thewing box 39 taken from above with the upper cover removed.FIG. 5 also showsribs 61, 62 (conventionally referred to as “rib two” and “rib three” respectively) which connect the fore and aft spars of thewing box 39 as well as the upper and lower covers. Theribs webs 46 of the fore and aft spars. -
FIG. 6 is similar toFIG. 4 and identical features are given the same reference numerals. Note that the lower cover has been omitted for clarity. In this case, the thickness of thecover 40 a is equal to the lower tolerance value X (seeFIG. 2 ). In order to achieve the required value of D, the thickness of thestack 52 a of thickness control plies is greater than the thickness of the (machined) stack 52 shown in Figure 4. When thecover 40 a and thestack 52 a of thickness control plies are bolted to the spar, the portion of the cover outboard of thestack 52 a is bent down towards theupper spar flange 44. This creates a joggledsection 70 a in thecover 40 a which is inclined in the opposite sense to theramp 55 in the stack of thickness control plies. - The angle of inclination of the ramp in the joggled
section 70 a is typically very shallow (i.e. within the bending allowables) and has little effect on the aerodynamic performance of the wing. Note that the angle of the joggledsection 70 a shown inFIG. 6 is exaggerated for clarity. -
FIG. 7 is similar toFIGS. 4 , 6 and identical features will be given the same reference numerals. Note that the lower cover has again been omitted for clarity. In this case the thickness of thecover 40 b is equal to the upper tolerance value Y (illustrated inFIG. 2 ). When thecover 40 b and thestack 52 b of thickness control plies are bolted to the spar, the portion of the cover outboard of thestack 52 b is bent up to create a joggledsection 70 b with a ramp which is inclined in the same sense as the ramp in the stack of thickness control plies. The ramp in the joggled cover engages the top of the ramp in the spar flange. - The ramp in the joggled
section 70 b is typically very shallow (i.e. within the bending allowables) and has little effect on the aerodynamic performance of the wing. Again, no shimming is required, which saves time and complexity. Note that, again, the angle of the joggledsection 70 b shown inFIG. 7 is exaggerated for clarity. Note also that the joggle has a downward inboard incline as opposed to the upward inboard incline of the joggledsection 70 a shown inFIG. 6 . - As noted above, in each case the
stacks stacks covers - In the method described above, the thickness of the cured panel assembly is first measured, and then if necessary some of the thickness control plies are removed from the
stack 26 by a machining process. In an alternative case in which the thickness control plies are bonded to a previously cured cover, then the thickness of the cured cover may be measured before the thickness control plies are bonded in place, and the number of thickness control plies selected in accordance with the measured thickness (either by peeling off some of the uncured plies in a pre-prepared stack, or by laying up the stack with exactly the right number of plies). Thus in this case it is not necessary to machine away some of the cured thickness control plies after they have been bonded to the cover. - Although the invention has been described above with reference to one or more preferred embodiments, it will be appreciated that various changes or modifications may be made without departing from the scope of the invention as defined in the appended claims.
Claims (15)
Priority Applications (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US13/944,489 US9162417B2 (en) | 2009-06-25 | 2013-07-17 | Method of manufacturing a structure |
Applications Claiming Priority (4)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
GB0910938.0 | 2009-06-25 | ||
GBGB0910938.0A GB0910938D0 (en) | 2009-06-25 | 2009-06-25 | Method of manufacturing a structure |
US12/821,718 US8524022B2 (en) | 2009-06-25 | 2010-06-23 | Method of manufacturing a structure |
US13/944,489 US9162417B2 (en) | 2009-06-25 | 2013-07-17 | Method of manufacturing a structure |
Related Parent Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US12/821,718 Division US8524022B2 (en) | 2009-06-25 | 2010-06-23 | Method of manufacturing a structure |
Publications (2)
Publication Number | Publication Date |
---|---|
US20140023830A1 true US20140023830A1 (en) | 2014-01-23 |
US9162417B2 US9162417B2 (en) | 2015-10-20 |
Family
ID=40972740
Family Applications (2)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US12/821,718 Expired - Fee Related US8524022B2 (en) | 2009-06-25 | 2010-06-23 | Method of manufacturing a structure |
US13/944,489 Active 2030-09-10 US9162417B2 (en) | 2009-06-25 | 2013-07-17 | Method of manufacturing a structure |
Family Applications Before (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US12/821,718 Expired - Fee Related US8524022B2 (en) | 2009-06-25 | 2010-06-23 | Method of manufacturing a structure |
Country Status (2)
Country | Link |
---|---|
US (2) | US8524022B2 (en) |
GB (1) | GB0910938D0 (en) |
Cited By (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20160121556A1 (en) * | 2009-12-10 | 2016-05-05 | Autodesk, Inc. | Composite part manufacturing compensation system and method |
US20170291320A1 (en) * | 2014-10-07 | 2017-10-12 | Ruairidh Robertson | Shaving Device |
Families Citing this family (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US8578609B2 (en) * | 2010-10-27 | 2013-11-12 | Airbus Operations S.L. | Manufacturing procedure of a part made out of a composite material introducing a geometric change in one of its faces |
ES2426111B1 (en) * | 2012-04-17 | 2015-03-24 | Airbus Operations S.L. | AIRCRAFT SUBSTITUTE SURFACE INTERFACE |
US10329009B2 (en) * | 2014-09-17 | 2019-06-25 | The Boeing Company | Composite wing edge attachment and method |
ES2887428T3 (en) * | 2016-11-16 | 2021-12-22 | Airbus Operations Gmbh | Airframe component for laminar flow |
CN109592005B (en) * | 2018-11-07 | 2021-10-15 | 中国航空工业集团公司西安飞机设计研究所 | Design method of composite material stringer with notch |
Citations (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20060188696A1 (en) * | 2005-02-24 | 2006-08-24 | The Boeing Company | Reinforced rampdown for composite structural member and method for same |
WO2008007043A1 (en) * | 2006-07-14 | 2008-01-17 | Airbus Uk Limited | Composite manufacturing method |
US20100136309A1 (en) * | 2007-04-13 | 2010-06-03 | Airbus France | Method for making a structural part of a composite material having a calibrated thickness and structural part thus obtained |
Family Cites Families (9)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US4565595A (en) * | 1981-09-30 | 1986-01-21 | The Boeing Company | Method of making composite aircraft wing |
US6749155B2 (en) * | 2002-09-13 | 2004-06-15 | The Boeing Company | Composite assembly with integrated composite joints |
JP4238311B2 (en) | 2002-12-19 | 2009-03-18 | 防衛省技術研究本部長 | 3D composite joint |
FR2866626B1 (en) * | 2004-02-20 | 2006-05-19 | Airbus France | OFFSET OF STIFF SLITTED SLOPES AND PANEL PROVIDED WITH SUCH A STOP |
DE102005034621B3 (en) | 2005-07-19 | 2007-01-11 | Deutsches Zentrum für Luft- und Raumfahrt e.V. | Composite structure of fiber-reinforced thermoplastics, e.g. slat with reinforcing rib, comprises first component welded to second component with endless fiber-reinforced and staple fiber-reinforced parts |
GB0525896D0 (en) * | 2005-12-20 | 2006-02-01 | Airbus Uk Ltd | A joint for use in aircraft construction |
ES2301359B1 (en) | 2006-05-16 | 2009-05-01 | Airbus España, S.L. | EXISTING HOLE SEALING SYSTEM BETWEEN THE FUSELAGE AND HEIGHT TIMON OF AN AIRCRAFT WITH ORIENTABLE HORIZONTAL STABILIZER. |
FR2911320B1 (en) | 2007-01-16 | 2009-04-10 | Sagem Defense Securite | AIRCRAFT VESSEL IN MULTIPLE DETACHABLE ELEMENTS |
US7712993B2 (en) | 2007-11-30 | 2010-05-11 | The Boeing Company | Double shear joint for bonding in structural applications |
-
2009
- 2009-06-25 GB GBGB0910938.0A patent/GB0910938D0/en not_active Ceased
-
2010
- 2010-06-23 US US12/821,718 patent/US8524022B2/en not_active Expired - Fee Related
-
2013
- 2013-07-17 US US13/944,489 patent/US9162417B2/en active Active
Patent Citations (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20060188696A1 (en) * | 2005-02-24 | 2006-08-24 | The Boeing Company | Reinforced rampdown for composite structural member and method for same |
WO2008007043A1 (en) * | 2006-07-14 | 2008-01-17 | Airbus Uk Limited | Composite manufacturing method |
US20100136309A1 (en) * | 2007-04-13 | 2010-06-03 | Airbus France | Method for making a structural part of a composite material having a calibrated thickness and structural part thus obtained |
Cited By (4)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20160121556A1 (en) * | 2009-12-10 | 2016-05-05 | Autodesk, Inc. | Composite part manufacturing compensation system and method |
US10286614B2 (en) * | 2009-12-10 | 2019-05-14 | Autodesk, Inc. | Composite part manufacturing compensation system and method |
US11167504B2 (en) | 2009-12-10 | 2021-11-09 | Autodesk, Inc. | Composite part manufacturing compensation system and method |
US20170291320A1 (en) * | 2014-10-07 | 2017-10-12 | Ruairidh Robertson | Shaving Device |
Also Published As
Publication number | Publication date |
---|---|
US8524022B2 (en) | 2013-09-03 |
US20100327114A1 (en) | 2010-12-30 |
GB0910938D0 (en) | 2009-08-05 |
US9162417B2 (en) | 2015-10-20 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
US9162417B2 (en) | Method of manufacturing a structure | |
EP3301014B1 (en) | Airfoil-shaped body having composite base skin with integral hat-shaped spar | |
EP2268539B1 (en) | Method of manufacturing a composite aircraft joint | |
EP2089218B1 (en) | Composite structure | |
US9669919B2 (en) | Highly integrated infused box made of composite material and method of manufacturing | |
US7806367B2 (en) | Integrated multispar torsion box of composite material | |
CN105438438B (en) | Composite wing edge attachment and method | |
EP2552678B1 (en) | Composite stringer and method of manufacturing a composite stringer | |
EP3287360B1 (en) | Aircraft composite wingbox integration | |
EP3287361B1 (en) | Planked stringers that provide structural support for an aircraft wing | |
WO2010144009A1 (en) | An aircraft structure with structural parts connected by nanostructure and a method for making said aircraft structure | |
US11673644B2 (en) | Composite spars with integrated sacrificial surfaces | |
US9868508B2 (en) | Rib foot for aircraft wing | |
US11577816B2 (en) | Composite laminate for an airframe lifting surface and method for manufacturing thereof | |
US11981426B2 (en) | Composite spars with integrated sacrificial surfaces | |
US20210122454A1 (en) | Composite laminate for an airframe lifting surface and method for manufacturing thereof |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
FEPP | Fee payment procedure |
Free format text: PAYOR NUMBER ASSIGNED (ORIGINAL EVENT CODE: ASPN); ENTITY STATUS OF PATENT OWNER: LARGE ENTITY |
|
AS | Assignment |
Owner name: AIRBUS OPERATIONS LIMITED, UNITED KINGDOM Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:ILIOPOULOS, ALEX;FLOOD, JOHN;SIGNING DATES FROM 20100906 TO 20101006;REEL/FRAME:036587/0152 |
|
STCF | Information on status: patent grant |
Free format text: PATENTED CASE |
|
MAFP | Maintenance fee payment |
Free format text: PAYMENT OF MAINTENANCE FEE, 4TH YEAR, LARGE ENTITY (ORIGINAL EVENT CODE: M1551); ENTITY STATUS OF PATENT OWNER: LARGE ENTITY Year of fee payment: 4 |
|
MAFP | Maintenance fee payment |
Free format text: PAYMENT OF MAINTENANCE FEE, 8TH YEAR, LARGE ENTITY (ORIGINAL EVENT CODE: M1552); ENTITY STATUS OF PATENT OWNER: LARGE ENTITY Year of fee payment: 8 |